The present invention lies in the field of electric thrust engines.
It applies in preferred but non-limiting manner to an ion or plasma thrust engine of the type used for delivering electric thrust in space, in particular for use with geostationary telecommunications satellites.
FIG. 1 is a general view of a prior art Hall effect plasma thruster 10. A central magnetic coil 12 surrounds a central core 14 that extends along a main longitudinal axis A. An annular inner wall 16 surrounds the central coil 12. This inner wall 16 is surrounded by an outer annular wall 18, the annular walls 16 and 18 defining an annular discharge channel 20 that extends around the main axis A. In the presently-described example, the inner wall 16 and the outer wall 18 form portions of a single ceramic part 19.
In the description below, the term “inner” designates a portion close to the main axis A and the term “outer” a portion remote from said axis.
Likewise, the terms “upstream” and “downstream” are defined relative to the normal flow direction (from upstream to downstream) of gas through the discharge channel 20.
The upstream end 20a of the discharge channel 20 (on the left in FIG. 1) is closed by an injection system 22 made up of a feed pipe 24 for feeding an ionizable gas (generally xenon), the pipe 24 being connected via a feed hole 25 to an anode 26 serving as a manifold for injecting gas molecules into the discharge channel 20.
The downstream end 20b of the discharge channel 20 is open (on the right in FIG. 1).
A plurality of peripheral magnetic coils 30, each presenting an axis parallel to the main axis A, are arranged all around the outer wall 18. The central magnetic coil 12 and the outer magnetic coils serve to generate a radial magnetic field B of intensity that is at a maximum at the downstream end 20b of the discharge channel 20.
A hollow cathode 40 is arranged outside the peripheral coils 30 so as to be oriented in order to eject electrons in a direction parallel to the main axis A and in the zone situated downstream from the downstream end 20b of the discharge channel 20. A potential difference is established between the cathode 40 and the anode 26.
The electrons as ejected in this way are directed in part into the inside of the discharge channel 20. Under the influence of the magnetic field generated between the cathode 20 and the anode 26, some of these electrons reach the anode 26, while most of them are trapped by the intense magnetic field B in the vicinity of the downstream end 20b of the discharge channel 20.
The gas molecules traveling from upstream to downstream in the discharge channel 20 are ionized by the electrons with which they come into collision.
Furthermore, the electrons present in the discharge channel 20 create an axial electric field E that accelerates the ions between the anode 26 and the downstream outlet 20b of the discharge channel 20 in such a manner that these ions are ejected at high speed from the discharge channel 20, thereby generating the thrust of the engine.
The invention relates more particularly to the feed system of the electric thruster.
In preliminary manner, it should be observed that present thrusters require a low regulated flow of gas in order to obtain thrust that is constant. This flow is created from a tank in association with a pressure regulator that brings the pressure into a constant range, the flow rate then being regulated so as to deliver the necessary quantity of gas to the engine and to the hollow cathode. This regulation is usually performed by a thermocapillary mechanism fed with electricity and by a flow rate restrictor enabling the flow to be shared between the anode and the cathode.
FIG. 2 shows a feed system 50 for the electric thruster 10 in accordance with the prior art.
That feed system 50 comprises a high-pressure tank 1 of ionizable gas, e.g. xenon or krypton, that is connected by a pipe 51 to a low-pressure buffer tank 2.
The volume of the low-pressure buffer tank 2 is about 1 liter (L).
The pressure in the high-pressure tank 1 varies from about 150 bars to about 1 bar; the pressure in the low-pressure buffer tank 2 varies in the range about 1.5 bars to 3 bars.
A restrictor 7 is placed in the pipe 51 to reduce pressure between the high-pressure tank 1 and the low-pressure buffer tank 2.
The pipe 51 also includes a regulator valve 6 for regulating the flow rate of gas between the high-pressure tank 1 and the low-pressure buffer tank 2.
The feed system 50 has means 53 for controlling the opening and closing of the regulator valve 6 and for measuring the pressure in the low-pressure buffer tank 2 in co-operation with a pressure sensor 54.
Downstream from the low-pressure buffer tank 2, the feed system 50 has two stop valves V3 and V4, a redundant stop valve V1, and a thermocapillary mechanism 52 for providing fine adjustment of the flow rate of gas to the anode 26 and to the cathode 40, respectively.
Restrictors 3 and 4 that are respectively associated with the anode 26 and with the cathode 40 serve to share the flow of gas between the cathode and the anode, with about 8% to 10% going to the cathode and about 90% to 92% going to the anode.
The feed system 50 also includes power electronics 81 suitable for applying voltage to the engine and ignition electronics 82 suitable for establishing a discharge current between the anode 26 and the cathode 40. Controlling software serves to sequence ignition of the engine and control of the vales for delivering gas and electricity to the thruster in a determined sequence.
In FIG. 2, discharge ignition as required solely for starting is referenced DA, and established engine discharge between the anode 26 and the cathode 40 is referenced DM.
It should be observed that with a Hall effect plasma engine, the above-mentioned electronics 81, 82 is often remote from the thruster, with a filter unit being used between the engine and the power electronics in order to avoid electromagnetic disturbances.
Usually, the subsystem constituted by the regulator valve 6, the restrictor 7, the low-pressure buffer tank 2, the means 53 for controlling opening and closing of the regulator valve 6, and the pressure sensor 54 constitutes a pressure regulator unit PRG.
Likewise, the stop valve V1, the thermocapillary mechanism 52, the restrictors 3, 4, and the valves V3, V4 constitute a unit RDX for regulating the flow rate of ionizing gas.
The thruster and the ignition system as described above present certain drawbacks.
Firstly, the bulk associated with the volume of the low-pressure buffer tank 2, typically 1 L, requires it to be offset in the satellite, thus needing additional tubular connections in the satellite. This arrangement is shown diagrammatically in FIG. 3 in which a satellite SAT presents a tubular connection between the pressure regulator unit PRG and the unit RDX for regulating the flow rate of ionizing gas.
Secondly, the feed system 50 requires the presence of valves downstream from the buffer tank (valves of the type V1, V3, and V4) in order to avoid losing the gas stored in the low-pressure buffer tank 2 when the engine is stopped, these valves being closed in practice simultaneously or practically simultaneously with the interruption of the supply of power to the engine.